U.S. patent application number 13/668274 was filed with the patent office on 2013-03-07 for transitional region for a secondary combustion chamber of a gas turbine.
This patent application is currently assigned to ALSTOM TECHNOLOGY LTD. The applicant listed for this patent is ALSTOM Technology Ltd. Invention is credited to Thomas HEINZ-SCHWARZMAIER, Paul MARLOW, Marc WIDMER, Selma ZAHIROVIC.
Application Number | 20130055717 13/668274 |
Document ID | / |
Family ID | 42340790 |
Filed Date | 2013-03-07 |
United States Patent
Application |
20130055717 |
Kind Code |
A1 |
HEINZ-SCHWARZMAIER; Thomas ;
et al. |
March 7, 2013 |
TRANSITIONAL REGION FOR A SECONDARY COMBUSTION CHAMBER OF A GAS
TURBINE
Abstract
A gas turbine is provided having a secondary combustion chamber
and a first guide vane row of a low-pressure turbine, the row being
arranged directly downstream of the chamber. The radially outer
boundary of the secondary combustion chamber is formed by at least
one outer wall segment, which is secured on at least one support
element arranged radially outwardly. The flow path of the hot gases
is bounded radially outwardly, in the region of the guide vane row,
by an outer platform which is secured at least indirectly on at
least one guide vane support. A substantially radially extending
gap-shaped cavity having a width in the range of 1-25 mm in the
axial direction in the inlet region is arranged between the wall
segment and the outer platform. At least one step element, which
reduces the width by at least 10% in at least one step, extending
substantially perpendicularly to the direction of flow of the hot
gas in the cavity, is arranged in the inlet region.
Inventors: |
HEINZ-SCHWARZMAIER; Thomas;
(Wettingen, CH) ; WIDMER; Marc; (Winterthur,
CH) ; ZAHIROVIC; Selma; (Basel, CH) ; MARLOW;
Paul; (Baden, CH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ALSTOM Technology Ltd; |
Baden |
|
CH |
|
|
Assignee: |
ALSTOM TECHNOLOGY LTD
Baden
CH
|
Family ID: |
42340790 |
Appl. No.: |
13/668274 |
Filed: |
November 4, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
PCT/EP2011/056582 |
Apr 26, 2011 |
|
|
|
13668274 |
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Current U.S.
Class: |
60/722 |
Current CPC
Class: |
F02C 6/003 20130101;
F01D 9/023 20130101; F01D 11/005 20130101; F05D 2240/127
20130101 |
Class at
Publication: |
60/722 |
International
Class: |
F02C 1/00 20060101
F02C001/00 |
Foreign Application Data
Date |
Code |
Application Number |
May 5, 2010 |
CH |
00691/10 |
Claims
1. A gas turbine having a secondary combustion chamber and a first
guide vane row of a low-pressure turbine, said row being arranged
directly downstream of said chamber, wherein a radially outer
boundary of the secondary combustion chamber is formed by at least
one outer wall segment, which is secured on at least one support
element arranged radially outwardly, wherein a hot gases flow path
is bounded radially outwardly, in a region of the first guide vane
row, by an outer platform which is secured at least indirectly on
at least one guide vane support, and wherein there is a
substantially radially extending gap-shaped cavity having a width
in the range of 1-25 mm in an axial direction in an inlet region,
between the at least one outer wall segment and the outer platform,
at least one step element, which reduces said width by at least 10%
in at least one step, extending substantially perpendicularly to
the direction of the hot gas flow in the cavity, is arranged in the
inlet region.
2. The gas turbine as claimed in claim 1, wherein the step element
is configured to encircle the axis of the turbine.
3. The gas turbine as claimed in claim 1, wherein the step element
is configured as encircling segments, and one such segment is
assigned radially outwardly to each guide vane, wherein all the
segments have a length in a circumferential direction, based on a
circumferential spacing of the guide vanes, of 30-50% of the
circumferential spacing.
4. The gas turbine as claimed in claim 1, wherein the step element
is in the form of a rib which is mounted or formed on the wall
region of the outer platform that adjoins the cavity, is
substantially rectangular in axial cross section and has a length
in the radial direction in the range of 10-100 mm.
5. The gas turbine as claimed in claim 1, wherein the step element
is in the form of a rib which is mounted or formed on a wall region
of the outer platform that adjoins the cavity, is substantially
rectangular in axial cross section and has a length in the radial
direction in the range of 20-50 mm, and is formed in combination
with a recess of equal or greater length, which is arranged
radially outwardly, which is formed in said wall region and the
radially outer end of which is formed by a further step.
6. The gas turbine as claimed in claim 5, wherein the outer
platform is secured on the guide vane support by an intermediate
ring, and wherein a further wall region of the cavity, said further
wall region radially adjoining the wall region of the outer
platform, is formed by said intermediate ring, and wherein a
further step is formed at the transition between the wall region of
the platform and the further wall region of the intermediate
ring.
7. The gas turbine as claimed in claim 1, wherein the cavity also
extends between the guide vane support and the support element.
8. The gas turbine as claimed in claim 1, wherein the step reduces
said width by at least 20%.
9. The gas turbine as claimed in claim 1, wherein the step reduces
said width by at least 30%.
10. The gas turbine as claimed in claim 5, wherein at least one
step element is arranged both on the outer platform and on the wall
of the wall segment and/or the wall region.
11. The gas turbine as claimed in claim 5, wherein a step element
is arranged only on the wall region of the outer platform and not
on the opposite wall of the wall segment, said wall being formed as
a radially extending plane.
12. The gas turbine as claimed in claim 1, wherein the radially
outwardly width of the step element increases substantially to the
original width in the inlet region, preferably via a step extending
substantially perpendicularly to the direction of flow of the hot
gas in the cavity.
13. The gas turbine as claimed in claim 1, wherein the width in the
inlet region is in a range of 2-15 mm in the axial direction.
14. The gas turbine as claimed in claim 1, wherein an encircling
projection, which locally narrows the inlet gap, is formed directly
at the inlet gap leading to the cavity, on the wall of the outer
wall segment.
Description
INCORPORATION BY REFERENCE
[0001] The following documents are incorporated herein by reference
as if fully set forth: International Patent Application No.
PCT/EP2011/056582, filed Apr. 26, 2011--and--Swiss Patent
Application No. 00691/10, filed May 5, 2010.
FIELD OF INVENTION
[0002] The present invention relates to a gas turbine, in
particular a transitional region between a secondary combustion
chamber and a low-pressure turbine in a gas turbine.
BACKGROUND
[0003] Gas turbines can be provided with a single combustion
chamber, but they can also have what is termed as sequential
combustion. In the case of the latter, fuel is burned in a first
combustion chamber, and the combustion air is then expanded via a
first turbine, a high-pressure turbine. Downstream of the
high-pressure turbine, the still hot combustion gases flow through
a secondary combustion chamber, in which more fuel is fed in and
typically burned in a process involving self-ignition. Arranged
downstream of this secondary combustion chamber is a low-pressure
turbine, by means of which the combustion gases are expanded, if
appropriate followed by a heat recovery system with steam
generation.
[0004] The transition of the housing from a combustion chamber to a
turbine is a critical region here because the temperature and
pressure conditions are particularly complex in this region.
Typically, the secondary combustion chamber, which is normally
designed as an annular combustion chamber, has, as it were, a
shell-shaped outer boundary, an outer wall which is composed of a
heat-resistant material or is correspondingly coated and which is
normally constructed from individual segments. On the opposite,
inner side, which is closer to the axis, there is a correspondingly
designed inner boundary, an inner wall composed of corresponding
materials. The low-pressure turbine, for its part, has a
multiplicity of alternately arranged rows of guide vanes and rotor
blades. The first row, which is arranged directly downstream of the
secondary combustion chamber, is typically a guide vane row
exhibiting a considerable twist of the vanes relative to the
direction of the principal axis. In this case, the guide vanes are
typically designed as segment modules, in which each guide vane has
an inner platform on the inside and an outer platform on the
outside, and the inner surfaces of these platforms then also form
the radially inner and radially outer boundaries of the flow
channel for the combustion air.
[0005] Accordingly, there is a gap on the radially inner side of
the annular flow channel between the inner wall segment of the
secondary combustion chamber and the inner platform of the first
guide vane row, and a gap on the radially outer side between the
outer wall segment of the secondary combustion chamber and the
outer platform of the first guide vane row. For reasons of assembly
and owing to the different mechanical and thermal loads on the
components comprising the secondary combustion chamber and the
turbine, this gap must have a certain width and cannot simply be
closed or fully bridged. The problem with this gap, which forms a
cavity that extends quite a long way radially towards the outside
into other structural components of the housing, especially on the
radially outer side, is the fact that it is furthermore exposed to
complex flow conditions, especially in the region of each guide
vane. This is because what is termed a bow wave or a "horse shoe
vortex" is formed at the leading edge of the guide vanes, leading
to hot combustion air being forced into this cavity in the wall
region and penetrating to a corresponding depth into the latter.
This can give rise to problems in connection not only with
overheating but also with oxidation of the corresponding
surfaces.
[0006] US 2009/0293488 discloses the possibility of substantially
closing this transitional region by means of a very small gap
dimension and additionally of providing specific structures which
ensure optimum cooling of the wall regions in this region. However,
the problem with this approach is that the required clearance
between the combustion chamber module and the turbine is not
automatically ensured as well, owing to the correspondingly small
gap dimension.
SUMMARY
[0007] The present disclosure is directed to a gas turbine having a
secondary combustion chamber and a first guide vane row of a
low-pressure turbine, the row being arranged directly downstream of
said chamber. A radially outer boundary of the secondary combustion
chamber is formed by at least one outer wall segment, which is
secured on at least one support element arranged radially
outwardly. A hot gases flow path is bounded radially outwardly, in
a region of the first guide vane row, by an outer platform which is
secured at least indirectly on at least one guide vane support. A
gap-shaped cavity having a width in the range of 1-25 mm in an
axial direction in an inlet region, extends substantially radially
between the at least one outer wall segment and the outer platform.
At least one step element, which reduces said width by at least 10%
in at least one step, extending substantially perpendicularly to
the direction of the hot gas flow in the cavity, is arranged in the
inlet region.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The following detailed description of the preferred
embodiment of the present invention will be better understood when
read in conjunction with the appended drawings. For the purpose of
illustrating the invention, there are shown in the drawings
embodiments which are presently preferred. It is understood,
however, that the invention is not limited to the precise
arrangements and instrumentalities shown. In the drawings:
[0009] FIGS. 1a-1d in: in 1a), an axial section of the transitional
region between the radial outer wall of the secondary combustion
chamber and the outer platform of the first guide vane row of the
low-pressure turbine, although the corresponding guide vane is not
shown, in 1b), a detail view of the section shown in a) with
illustrated hot air flows in the cavity, in 1c), a contour
illustration of the cavity and, in 1d), a schematic illustration of
the flow conditions in the inlet region of the cavity;
[0010] FIGS. 2a-2c show: in 2a), a detail view of a cavity with a
step element, in 2b), a contour illustration of a cavity with a
step element and, behind the latter, a set-back wall of the outer
platform and, in 2c), a schematic illustration of the flow
conditions in the inlet region of the cavity with a step element;
and
[0011] FIGS. 3a-3b show, in 3a), a schematic view, in a radial
direction, of a segment of the cavity with an encircling step
element and, in 3b), a corresponding view with a row of step
element segments.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Introduction to the Embodiments
[0012] It is here that the present invention intervenes, taking a
completely different approach to that of the prior art. More
specifically, there will be no attempt to close the gap since this
results in the abovementioned problems. On the contrary, although
the gap will have a certain width (in the axial direction),
appropriate measures will be taken to ensure that hot air or
combustion air is prevented from readily entering said gap and
potentially causing the abovementioned problems.
[0013] To be specific, the present invention accordingly relates to
a gas turbine having a secondary combustion chamber and a first
guide vane row of a low-pressure turbine, said row being arranged
directly downstream of said chamber, wherein the radially outer
boundary of the secondary combustion chamber is formed by at least
one outer wall segment, which is secured on at least one support
element arranged radially on the outside, wherein the flow path of
the hot gases is bounded radially on the outside, in the region of
the guide vane row, by an outer platform which is secured at least
indirectly on at least one guide vane support, and wherein there is
a substantially radially extending gap-shaped cavity having a width
B in the range of 1-25 mm in the axial direction in the inlet
region, between the wall segment and the outer platform. The width
B is specified for the cold installation condition. Depending on
the size of the housing clearances and permitted tolerances, the
width B is in a range of 2-15 mm.
[0014] According to the invention, this gap is not closed or
narrowed to an extreme extent as regards the gap dimension, at
least in the inlet region; rather, the approach followed is such
that at least one step element, which reduces said width B by at
least 10% in at least one step extending substantially
perpendicularly to the direction of flow of the hot gas in the
cavity, is arranged in the inlet region.
[0015] This step element, which is arranged substantially directly
behind the actual inlet region (typically 10-50 mm radially to the
outside of the inlet gap), produces flow vortices which to a
certain extent assume a sealing function and prevent the hot air
from penetrating deep into the cavity. Thus, it is also important
that the step should be able to produce such turbulence, and
therefore the step is preferably designed as a single step which
achieves the specified reduction of at least 10% in a single stage.
Typically, the step has substantially right-angled transitional
regions.
[0016] According to a first preferred embodiment, the step element
is designed to encircle the axis of the turbine. Accordingly, the
step element is designed substantially as an encircling rib
arranged in the cavity, on one of the walls of the latter. It is
possible for a single such step element to be arranged in the
cavity but it is also possible for a plurality of such step
elements to be provided in a manner radially offset towards the
outside. Accordingly, it is possible to widen the cavity again
behind the first step and to provide a second step element after
said widening. Thus, two vortices are produced and an enhanced
sealing function is ensured. If the width B of the cavity is
sufficient, at least one further step element can be arranged on
the wall of the cavity opposite the first step element. Typically,
the step elements are situated opposite one another, giving rise to
a constriction from both sides of the cavity.
[0017] Problems arise, in particular, in those regions which are
arranged radially directly to the outside of the respective leading
edge of each guide vane. In these regions, in particular, the
combustion air made turbulent by the bow wave has a particularly
strong tendency to penetrate into the cavity. The intermediate
regions, in contrast, are less strongly affected. Accordingly, it
is also possible, according to another preferred embodiment, for
the step element to be designed as encircling segments, with one
such segment being assigned radially on the outside to each guide
vane (that is to say regions of the cavity that are situated
between the segments do not have a step element). It is preferable
if substantially all the segments have a length in the
circumferential direction, based on the circumferential spacing p
(pitch) of the guide vanes, of 30-50% of the circumferential
spacing p.
[0018] By way of example, the step elements, designed as encircling
segments, can be assigned symmetrically to the guide vanes (that is
to say extending circumferentially by the same amount in the
clockwise direction and in the counterclockwise direction from the
radial position of the leading edge) or can be arranged offset with
respect to the guide vanes in a manner corresponding to a radial
offset of the bow wave.
[0019] Another preferred embodiment of the gas turbine proposed is
characterized in that the step element is in the form of a rib
which is mounted or formed on the wall region of the outer platform
that adjoins the cavity and is substantially rectangular in axial
cross section. Preferably, the rib has a length in the radial
direction in the range of 10-100 mm, particularly preferably in the
range of 20-50 mm. It is furthermore preferred if the rib is used
in combination with a recess of equal or greater length, which is
arranged radially on the outside, which is formed in this wall
region and the radially outer end of which is formed by a further
step, giving rise, radially in series, to two or three vortices and
ensuring an enhanced sealing effect.
[0020] In general terms, it is preferred if the wall, which is
situated opposite the step element, bounds the cavity and extends
substantially perpendicularly to the axis of the turbine does not
itself have a step element. In other words, in the present
invention, it is preferentially not a matter of providing a
labyrinth seal in the traditional sense, in which the flow path is
as it were designed in a meandering shape; rather the point is to
provide a step element on only one of the two opposite walls of the
cavity. In fact, labyrinth seals can be problematic since they can
restrict the clearance function of the gap and have a negative
impact on ease of assembly.
[0021] In general terms, the step element or the plurality of
segments, in which one step element is assigned to each guide vane,
is preferably arranged on the wall situated downstream in the
direction of flow of the hot gas in the secondary combustion
chamber, i.e. normally on the platform.
[0022] According to another preferred embodiment, the outer
platform is secured on the guide vane support by means of an
intermediate ring, wherein a further wall region of the cavity,
said wall region radially adjoining the wall region of the outer
platform, is formed by this intermediate ring, and wherein
furthermore a further step is formed, preferably at the transition
between the wall region of the platform and the further wall region
of the intermediate ring.
[0023] The cavity preferably also extends between the guide vane
support and the support element, i.e. it is a cavity which extends
deep into the structure.
[0024] According to another preferred embodiment of the invention,
said width B is reduced by at least 20%, preferably by at least
30%, by the step (designed as a single step). Under specific
conditions, a reduction by at least 40% may even be desirable.
Typically, a reduction by up to 70% is desirable. Any reduction
beyond this is generally not practicable and could furthermore have
an effect on desired purging flows.
[0025] As already mentioned above, it is preferred if a step
element is arranged only on the wall region of the outer platform
and none on the opposite wall of the wall segment, said wall
preferably being formed as a radially extending plane. An outer
platform of this kind does not necessarily extend a long way
radially towards the outside. In this case, this wall region on
which the step element is arranged is then also not formed by the
platform but is formed by the intermediate ring arranged to the
outside or by the guide vane support.
[0026] It is preferable if the width of the cavity radially to the
outside of the step element increases again substantially to the
original width B in the inlet region, preferably via a step
extending substantially perpendicularly to the direction of flow of
the hot gas in the cavity, and, as a further preference, said step
is followed in a radially outward direction by a second step, which
once again narrows.
[0027] The width B in the inlet region is preferably in the range
of 1-25 mm in the axial direction.
[0028] It is possible that an encircling projection, which locally
narrows the inlet gap, is formed directly at the inlet gap leading
to the cavity, on the wall of the outer wall segment.
DETAILED DESCRIPTION
[0029] FIG. 1a shows an axial section through the radially outer
wall region of a gas turbine having a secondary combustion chamber
1, at the transition from the secondary combustion chamber 1 to the
first guide vane row 2 of the low-pressure turbine. The radially
inner boundary of the flow channel for the hot gases 3 is not
shown. Radially on the outside, the flow channel within the
secondary combustion chamber 1 is formed by an outer wall segment
4. This is typically composed of metal or ceramic, and the metal is
typically provided with a thermal protective coating. This outer
wall segment 4 is secured on the housing by means of a support
element 5 and is normally supplied at the rear with appropriate
cooling air flows, which may additionally emerge into the hot air
flow through cooling air openings in the wall segment 4 to give
film cooling.
[0030] Downstream in the direction of flow of the hot gas 3, the
secondary combustion chamber is followed by the first guide vane
row 2. Guide vanes are typically integral structures which comprise
not only the actual guide vane but also an inner platform and an
outer platform 6 integrally formed thereon. The guide vanes can
also be grouped into subassemblies comprising a plurality of guide
vanes. The platforms, which cover a segment when viewed in a
direction around the turbine axis, not only form the fastening of
each of the guide vanes when a row of such guide vane elements is
arranged around the circumference of a gas turbine, but
simultaneously also form the radially outer boundary of the flow
path for the hot gas in the case of the outer platform 6 and the
inner boundary of said flow path in the case of the inner platform.
In other words, the outer platforms 6 form an encircling ring which
tapers in the direction of flow. The outer platforms 6 or said
units of guide vanes and inner and outer platforms 6 are secured on
what is termed an intermediate ring 7, which, for its part, is
secured on the housing on what is termed a guide vane support 8 of
the low-pressure turbine.
[0031] A gap is formed between the wall elements 4 of the secondary
combustion chamber 1 and the outer platform 6 of the first guide
vane row 2 of the low-pressure turbine, said gap forming a cavity 9
that extends deep into the housing components.
[0032] This cavity 9 is shown in greater detail in FIG. 1b. Owing
to the bow wave, already described at the outset, at the leading
edge of each guide vane, there is a high hot gas pressure in the
inlet region of said cavity 9, especially at these radial
positions. Accordingly, there is a hot gas flow, indicated
schematically by the arrow 10, into this inlet region, penetrating
deep into the cavity, as illustrated schematically by arrow 11.
Here, the cavity 9 is initially formed on the downstream side
(relative to the main direction of flow of the hot gases 3), by a
wall region 12 of the outer platform 6, followed by a wall region
13 of the intermediate ring 7 and, further towards the outside
radially, by a wall region 14 of the guide vane support 8. In the
prior art design, these wall regions 12-14 lie substantially flush
in one plane. The boundary wall of the cavity 9 which is arranged
opposite and further upstream in the direction of flow is initially
formed, radially on the inside, by the wall region 15 of the outer
wall segment 4 of the secondary combustion chamber, followed,
radially on the outside, by the wall region 16 of the support
element 5 for the wall segment 4. Here as well, these wall regions
15, 16 are flush in the prior art designs. The hot air flow 11 not
only has the effect that unnecessarily high temperatures are
reached in the cavity but also leads, in particular, to oxidation
problems in wall regions 12-16. On the other hand, this gap is
necessary for assembly reasons.
[0033] In the inlet region 27, this gap or cavity 9 has a width B,
which is indicated in the contour illustration of the cavity 9 in
FIG. 1c. This width is typically in the range of 1-25 mm, i.e. the
gap is wide in this region and correspondingly accessible for said
hot air flow. Directly at the inlet gap 17 leading into this cavity
9, there is an encircling projection 18 extending in the direction
of flow of the hot gas from the outer wall segment 4, on the
radially forward edge of the latter, said projection reducing the
inlet or front inlet gap width somewhat. Behind this, however, the
inlet gap widens again to said width B.
[0034] In the case of such a gap without special measures, the flow
pattern which forms is as illustrated schematically in FIG. 1d. The
hot gas passes through the inlet gap 17 and past the encircling
projection 18 and forms a hot air vortex 20 behind said projection
in the inlet region. Radially to the outside of this vortex, the
hot gas then flows substantially unhindered in a radial direction
and, accordingly, flows at high temperatures, i.e. with a high
oxidative effect, deep into the gap of the cavity 9.
[0035] FIG. 2a shows a detail similar to FIG. 1b, which is
additionally formed with a step element 22 according to the
invention. This step element is designed as an encircling rib,
which is arranged on the wall region 12 or is formed integrally
with the latter and provides a radially inner step immediately
downstream of the encircling projection 18 in the direction of flow
of the hot gas 10. Typically, this step element 22 extends in a
radial direction approximately over one third or even half the
radial extent of the wall region 12. Apart from the encircling
projection 18 in the inlet gap 17, the opposite wall 15, by
contrast, is of flat configuration and is not likewise formed with
a step element or with an appropriately corresponding groove.
Accordingly, the step element 22 to a certain extent forms a
barrier for the hot gas flow, and turbulence reduces the speed of
the hot gas. Accordingly, leakage flows and purge air flows can
then cool and protect the corresponding wall regions in a
significantly more efficient manner. Both the step of the step
element 22 which faces the inlet gap and the radially outer step
behind the step element 22, where the cavity widens again, lead to
vortex formation.
[0036] In the contour illustration shown in FIG. 2b, not only is
the step element 22 additionally formed on the outer platform 6 but
the wall region behind it is also cut out or recessed somewhat,
with the result that the width is increased somewhat more than
previously radially to the outside of the step element 22, and a
pronounced step 29 is then also formed at the transition 23 to the
wall region 13. This step 29 leads to additional turbulence and an
expanded additional barrier function.
[0037] FIG. 2c illustrates schematically the flow conditions with
such a construction. As before, there is a first vortex 20
essentially behind the encircling projection 18, but this is
significantly intensified by the inlet step of the step element 22.
In other words, this vortex is significantly more powerful than in
FIG. 1 and also develops a greater barrier effect. In addition, a
first vortex 24 is formed in the region of the step element 22. A
second vortex 25 is formed to a certain extent at the radially
outer end of the step element in the region where it widens, and
these vortices 24, 25 lead to an additional barrier effect.
Depending on the detailed geometry and purge air flow, an
additional step 29 at the transition 23 promotes turbulence and
leads to a further additional barrier function. If the temperatures
are now observed, it will be ascertained that the temperature can
be extremely reduced by these measures, not just in the region of
the step element 22 but also radially to the outside thereof, with
the result that lower pressures prevail and, accordingly, the
regions arranged in the region of the step element 22 and radially
to the outside thereof can be protected significantly more easily
with cooling air.
[0038] FIG. 3a illustrates how the step element 22' can be of
encircling design, i.e. in the form of a substantially encircling
ring around the axis of the low-pressure turbine. As already
explained at the outset, those problems which are actually serious
occur mainly at the leading edge of the respective guide vane 26.
Accordingly, it may also be sufficient, as illustrated in FIG. 3b,
if only a segment 22'' of such a step element is arranged as it
were radially to the outside of each guide vane and in a manner
coordinated with the leading edge thereof in order to produce the
effect according to the invention.
[0039] It is understood, therefore, that this invention is not
limited to the particular embodiments disclosed, but is intended to
cover all modifications which are within the spirit and scope of
the invention as defined by the appended claims; the above
description; and/or shown in the attached drawings.
LIST OF REFERENCE SIGNS
[0040] 1 secondary combustion chamber [0041] 2 guide vane row
[0042] 3 hot gas flow [0043] 4 outer wall segment of 1 [0044] 5
support element for 4 [0045] 6 outer platform of 26 [0046] 7
intermediate ring [0047] 8 guide vane support of the low-pressure
turbine [0048] 9 outer cavity [0049] 10 hot gas flow inlet in 9
[0050] 11 hot gas flow in 9 [0051] 12 wall region of 6 adjoining 9
[0052] 13 wall region of 7 adjoining 9 [0053] 14 wall region of 8
adjoining 9 [0054] 15 wall region of 4 adjoining 9 [0055] 16 wall
region of 5 adjoining 9 [0056] 17 inlet gap in 9 [0057] 18
encircling projection [0058] 20 vortex in the inlet region [0059]
22 step element [0060] 22' step element, encircling [0061] 22''
step element, in segments [0062] 23 step transition from 12 to 13
[0063] 24 first vortex [0064] 25 second vortex [0065] 26 guide vane
[0066] 27 inlet region of 9 [0067] 28 first step on 22 [0068] 29
step at 23 [0069] p pitch [0070] B width in the inlet region
* * * * *